Gas turbine engines operate at elevated temperatures, and film cooling is widely used to protect components from the harsh high-temperature environment. Maintaining metal temperatures for gas turbine components within material limits has been addressed by many different techniques such as film cooling, impingement cooling, low conductivity coatings and heat augmentation devices such as turbulators, ribs, pin fin banks, etc.
Film cooling is widely used in connection with gas turbine first-stage components and to a lower extent in subsequent stages. Standard practice among the industry is to feed these film cooling holes from existing cavities built into the component. This severely limits flexibility with respect to drilling holes at locations not aligned with the cavities. As a result, the designer oftentimes cannot place film cooling at locations of high level temperatures, or has to orient the cooling holes at angles that reduce the impact of the film cooling. Competitors have addressed this issue in the past by machining dedicated chambers and serpentine passages into the component. These features are only manufactured for the purpose of feeding these holes, and add extra manufacturing cost to the component.
Specific examples in the prior art include cooling holes fed from cavities cast into the turbine sidewalls as exemplified by U.S. Pat. No. 5,344,283. Other approaches for casting dedicated chambers into the sidewalls with the intent of feeding film cooling holes are disclosed in U.S. Pat. Nos. 6,254,333 and 6,210,111. A cavity formed by seal plates in a cold side of a stage one turbine nozzle is disclosed in U.S. Pat. No. 5,417,545. A concept for machining multiple cooling holes such that they feed from the same aperture in a cold side cavity is disclosed in U.S. Pat. No. 5,062,768. The assignee of this invention presents a concept for pressurizing a seal slot with air from cooling cavities for the purpose of cooling the seal itself in U.S. Pat. No. 6,340,285.